Rotary turbine engine

ABSTRACT

Intake air undergoes two stage compression within a pair of coaxially mounted, counter-rotating rotors of a radial turbine engine and is tangentially discharged into a turbine mounted by one of the rotors. The air is heated and accelerated by discharge of combustion products from fuel burners mounted on the second stage compressor vanes in the other of the rotors to produce reaction thrust. Extensions of some of the second stage compressor vanes form discharge passages for the mixture of air and combustion products.

United States Patent 91 Fincher [451 Apr. 17, 1973 [5 ROTARY TURBINEENGINE 820,337 9/1959 Great Britain... Who/39.35 Inventor: Julian D. R OBox 801,281 9/1958 Great Britain 60/3935 Pensacola, Fla. 32502 PrimaryExammerCar1ton R. Croyle Flledl M811 1971 Assistant ExaminerWarren Olsen[52] US. Cl ...60/39.35, 60/39.16 C

[51] Int. Cl. ..F02c 3/16 [58] Field of Search ..60/39.35, 39.34, 60/39.16 C; 415/63 [56] References Cited UNITED STATES PATENTS 3 ,077,0752/1963 Turanciol ..60/39.35

1,868,143 7/1932 Heinze ..60/39.35

3,557,551 1/1971 Campbell ..60/39.35

1,003,708 9/1911 Coleman ..60/39.35

Appl. No.: 125,931

FORElGN PATENTS OR APPLICATIONS 1/1948 France ..60/39.35

Attorney-Clarence A. OBrien and Harvey B. Jacobson ' ABSTRACT lntake airundergoes two stage compression within a pair of co-axially mounted,counter-rotating rotors of a radial turbine engine and is tangentiallydischarged into a turbine mounted by one of the rotors. The air isheated and accelerated by discharge of combustion products from fuelburners mounted on the second stage compressor vanes in the other of therotors to produce reaction thrust. Extensions of some of the secondstage compressor vanes form discharge passages for the mixture of airand combustion products.

6 Claim, 6 Drawing Figures PATENTEDAPR I 71973 SHEET 1 UP 3 Julian 0.Fincher 1N1 'EXTOK 9 IS) WW 19% PATENTED APR 1 7 I973 SHEET 2 [1F 3Ju/ian 0. Fin char PAIENIED R 3,727,401

Jul/an 0. F l'ncher INVENTOI-i.

dye/way 19% ROTARY TURBINE ENGINE This invention relates to a radialturbine type of engine and more particularly to a unique structuralarrangement for such an engine. a

The present invention pertains to a compact gas turbine engine fromwhich power is obtained in the form of a rotating power shaft. Theengine features a pair of V counter-rotating rotors one of which isfixed to the power shaft. Both rotors mount compressor vanes for twostage compression of air. The air is drawn in through the compressorblades of one of the rotors on which the turbine blades are also mountedradially outwardly of the compressor blades on the other of the rotorsfixed to the power shaft. Thus, the air is sequentially compressedradially inwardly and then radially outwardly before being dischargedinto the turbine blades to impart rotation to the turbine rotor. Fuelburners on the power shaft rotor discharge jets of combustion productsso as to impart thermal energy and kinetic energy to the air duringcompression as well as to produce reaction thrust imparting rotation tothe power shaft mounted rotor. The mixture of air and combustionproducts exits from the turbine blades within an outer manifold housingfrom which the mixture is exhausted.

These together with other objects and advantages which will becomesubsequently apparent reside in the details of construction andoperation as more fully hereinafter described and claimed, referencebeing had to the accompanying drawings forming a part hereof, whereinlike numerals refer to like parts throughout, and in which:

FIG. 1 is a perspective view showing a typical gas turbine engineconstructed in accordance with the present invention.

FIG. 2 is a side sectional view taken substantially through a planeindicated by section line '2-2 in FIG. 1.

FIG. 3 is a transverse sectional view taken substantially through aplane indicated by section line 3-3 in FIG. 2.

' FIG. 4 is a side sectional view taken substantially through a planeindicated by section line 4--4 in FIG. 3

FIG. 5 is a perspective view showing one of the rotor assembliesassociated with the turbine engine.

FIG. 6 is a perspective view showing the other rotor assembly associatedwith the gas turbine engine with a portion broken away.

Referring now to the drawings in detail, FIG. 1 illustrates oneembodiment of a gas turbine engine constructed in accordance with thepresent invention and generally denoted by reference numeral 10. Theengine is mounted by a suitable frame generally referred to by referencenumeral 12 having a base portion 14 and a pair of standards 16 mountingbearing blocks 18 through which a power delivery shaft 20 extendsadapted to be connected to some load at one end by means of a key 22. Asuitable fuel is supplied to the engine through the other end of thepower delivery shaft 20'by means of the fuel supply line 24 and thefitting 26. The frame also includes a base block portion 28 to which amanifold housing 30 is fixedly secured. The manifold housing may beformed by two sections interconnected by a plurality of fasteners 32along abutting flange portions 34. The manifold housing may be of agenerally annular shape converging at one circumferential locationtoward an exhaust opening surrounded by a circular flange 36 to which anexhaust conduit is adapted to be secured along an axis perpendicular tothe rotational axis of the power delivery shaft 20.

As more clearly seen in FIG. 3, the power shaft is axially fixed betweenthe bearing blocks 18 by means of the thrust washer 38 and the set screwmounted collar 40 axially abutting sleeve bearings 42 journaling thepower shaft. Sleeve bearings 44 are also supported on the power shaftspaced from the bearings 42 by spacers 46 in order to rotatably supporta free-running rotor assembly generally referred to by reference numeral48 that is not mechanically loaded. A second rotor assembly 50 isaxially and rotationally fixed to the power shaft 20 within the rotorassembly 48.

Referring now to FIGS. 2, 3 and 6, the rotor assembly 50 includes a hubdisc 52 secured to the power shaft 20. Secured to the hub disc 52 andprojecting radially outwardly therefrom are a plurality of curvedcompressor vanes 54 that are relatively short in length as compared toalternatively spaced longer compressor vanes 56 provided withcircumferentially overlapping extensions 58 terminating at the radiallyouter periphery of a pair of circular end wall discs 60 and 62 betweenwhich the compressor vanes 54 and 56 are disposed and to which they areconnected as by welding. The compressor vanes furthermore convergeradially outwardly so as to form volumetrically decreasing flow areatherebetween and within the discharge passages 64 formed between theoverlapping portions of the vanes 56 and extensions 58. The end walldiscs 62 are provided with central inlet openings 66 into which anintake fluid such as air is admitted for flow radially outwardly duringrotation of the rotor assembly 50. Accordingly, fluid will be dischargedtangentially from the discharge passages 64 of the rotor assembly 50after undergoing compression.

The fluid discharged from the rotor assembly 50, impinges uponcircumferentially spaced turbine blades 68 fixed between the radiallyouter peripheral portions of a pair of intermediate discs 70 associatedwith the rotor assembly 48. With reference to FIGS. 2, 3, 4 and 6, theintermediate discs 70 of the rotor assembly 48 are interconnected withend wall discs 72 by first stage compressor vanes 74. The end wall discs72 are provided with sleeve portions 76 supported on the sleeve bearings44 aforementioned. Further, the end wall discs 72 are of a smallerdiameter than the intermediate discs 70. The intermediatedi scs 70 areformed with central outlet openings 78 aligned with the inlet openings66 in the rotor assembly 50 aforementioned. Further, the intermediatediscs 72 converge radially outwardly to enclose the rotor assembly 50therebetween and extend into the manifold housing 30 so as to positionthe turbine blades 68 radially outwardly of the rotor assembly 50.

As more clearly seen in FIG. 2, the longer compressor vanes'56 of rotorassembly 50, are provided with fuel passages 80 extending to fuelbumers82 through which the fuel is burned and combustion products dischargeinto the discharge passages 64. The fuel passages 80 extend from thefuel burners 82 through the vanes 56 and the hub disc 52 to portsterminating at the axially inner end of a fuel passage 84 formed in thepower shaft 20 as shown in FIG. 3. The fuel passage 80 communicatesthrough the fitting 26 with the fuel line 24.

It will be apparent from the foregoing description of the gas turbineengine, that once the fuel is ignited and combustion products aredischarged from the fuel burners 82, a reaction thrust is producedcausing rotation of the rotor assembly 50 in a clock-wise direction asviewed in FIG. 2. The fluid thus discharged from the discharge passages64 impinges on the turbine blades 68 producing rotation of theassociated rotor assembly 48 in an opposite or counterclock-wisedirection. The rotors 48 and 50 thus rotate in opposite directions toeffect two stage compression of the air. The compressor vanes 74associated with the rotor 48 initially compress the air drawn radiallyinwardly on either axial side of the exhaust manifold housing 30. Theinitially compressed air passes through the outlet openings 78 and intothe inlet openings 66 for second stage compression by the vanes 54 and56 associated with the rotor 50. While undergoing second stagecompression, the air is heated by the combustion products andfurthermore accelerated, thereby. Thus, a fluid mixture emerges from thedischarge passages 64 with increased kinetic energy and thermal energywhich is substantially absorbed by the turbine blades 68 for poweringthe first stage compressor blades 74. Of course, a major portion of theenergy released-by the burner is in the form of reaction thrustproducing rotation of the rotor 50 and the power delivery shaft 20connected thereto. Further, it will be apparent that only the rotor 50is mechanically loaded so that any increase in load causing a decreasein its speed will produce an increase in speed of rotor 48. This occursbecause the thrust of the gas discharged peripherally from rotor 50increases with any decrease in speed to increase the kinetic energyimparted to rotor 48 through turbine blades 68. The increased speed ofrotor 48 thereby increases the inflow rate of air drawn in by itsfirststage compressor blades 74 to more completely support fuelcombustion that may be increased to meet the increased load on the powershaft.

The foregoing is considered as illustrative only of the principles ofthe invention. Further, since numerous modifications and changes willreadily occur to those skilled in the art, it is not desired to limitthe invention to the exact construction and operation shown anddescribedQand accordingly all suitable modifications and equivalents maybe resorted to, falling within the scope of the invention.

What is claimed as new is as follows:

1. A gas turbine engine comprising a frame, a pair of counter-rotatingrotors rotatably mounted by the frame, one of said rotors beingfree-running, radial compressor means mounted by said rotors for twostage compression of an intake fluid, turbine blade means mounted bysaid one of the rotors for receiving the in take fluid discharged fromthe compressor means and combustion means mounted by the other of therotors for heating the, intake fluid during compression, said one of therotors including an intake section through which the compressor meansdraws the intake fluid radially inward and a radially outer sectionmounting the turbine blade means in enclosing relation to the other ofthe rotors.

2. The combination of claim 1 wherein the other of the rotors includes apower delivery member on which said one of the rotors is rotationallysupported, and a pair of end walls having inlet openings and projectingradially outward into the radially outer section of said one of therotors.

3. The combination of claim 2 wherein said compressor means includesfirst stage compressor vanes mounted by the intake section of said oneof the rotors, second stage compressor vanes mounted by the other of therotors radially inwardly of the turbine blade means, and extensions onsome of said second stage compressor vanes forming discharge passagesfor the intake fluid.

4. The combination of claim 3 wherein the combustion means includesthrust producing fuel burners mounted by the other of the rotors fromwhich jets of combustion products are discharged into the intake fluidconducted by said discharge passages.

5. The combination of claim 1 wherein said compressor means includesfirst stage compressor vanes mounted by the intake section of said oneof the rotors, second stage compressor vanes mounted by the other of therotors radially inwardly of the turbine blade means, and extensions onsome of said second stage compressor vanes forming discharge passagesfor the intake fluid.

6. The combination of claim 5 wherein the combustion means includesthrust producing fuel burners mounted by the other of the rotors fromwhich jets of combustion products are discharged into the intake fluidconducted by said discharge passages.

1. A gas turbine engine comprising a frame, a pair of counterrotatingrotors rotatably mounted by the frame, one of said rotors beingfree-running, radial compressor means mounted by said rotors for twostage compression of an intake fluid, turbine blade means mounted bysaid one of the rotors for receiving the intake fluid discharged fromthe compressor means and combustion means mounted by the other of therotors for heating the intake fluid during compression, said one of therotors including an intake section through which the compressor meansdraws the intake fluid radially inward and a radially outer sectionmounting the turbine blade means in enclosing relation to the other ofthe rotors.
 2. The combination of claim 1 wherein the other of therotors includes a power delivery member on which said one of the rotorsis rotationally supported, and a pair of end walls having inlet openingsand projecting radially outward into the radially outer section of saidone of the rotors.
 3. The combination of claim 2 wherein said compressormeans includes first stage compressor vanes mounted by the intakesection of said one of the rotors, second stage compressor vanes mountedby the other of the rotors radially inwardly of the turbine blade means,and extensions on some of said second stage compressor vanes formingdischarge passages for the intake fluid.
 4. The combination of claim 3wherein the combustion means includes thrust producing fuel burnersmounted by the other of the rotors from which jets of combustionproducts are discharged into the intake fluid conducted by saiddischarge passages.
 5. The combination of claim 1 wherein saidcompressor meAns includes first stage compressor vanes mounted by theintake section of said one of the rotors, second stage compressor vanesmounted by the other of the rotors radially inwardly of the turbineblade means, and extensions on some of said second stage compressorvanes forming discharge passages for the intake fluid.
 6. Thecombination of claim 5 wherein the combustion means includes thrustproducing fuel burners mounted by the other of the rotors from whichjets of combustion products are discharged into the intake fluidconducted by said discharge passages.